This invention relates to airfoils for aerospacecraft, and more particularly to a ruddervator for an aerospacecraft which comprises a hybrid construction that enables lower production costs and improved life and reliability of the ruddervator.
Current control surfaces for advanced aerospacecraft are formed by a carbon-based ceramic matrix composite (CMC) hot structure with conventional rib-stiffened structure and a mechanically fastened skin. The X-37 aerospacecraft presently in use incorporates a control surface termed a xe2x80x9cruddervatorxe2x80x9d with the above-described construction which makes use of carbon/silicon carbide (C/SiC). This construction is shown in FIG. 1. The mechanically fastened upper skin 10 is secured by a high temperature metal, ceramic or ceramic composite fasteners at locations 12 to an integral C/SiC lower skin and substructure 14. A C/SiC tail tip 16 is used to close the end of the ruddervator. A titanium spindle 17 is used to rotate the ruddervator as needed. Thermal protection system seals 18, 20 and ring 22 are used to help mount the ruddervator to the fuselage of the aerospacecraft.
The X-37 ruddervator approach described above uses an expensive 2800xc2x0 F. CMC system in a 2400xc2x0 F. xe2x80x9chot structurexe2x80x9d application and uses an aircraft-like structural approach at the elevated temperature. The term xe2x80x9chot structurexe2x80x9d refers to the temperature of the primary load-carrying structure, in this case the CMC and supports used at 2400xc2x0 F. This construction reduces the service life of the fasteners. Furthermore, carbon-based CMCs generally require complex and costly tooling, unique and expensive infiltration/furnace facilities, and fabrication cycles of six months or more. The use of new materials under development, such as oxide fibers/oxide matrix-based CMC (Oxide-CMC), provide opportunities to design control surfaces in more cost-effective ways including, but not limited to, maintaining internal supports and attachments below 600xc2x0 F.
For present and planned reusable hypersonic vehicles there are also size constraints on control surfaces due to available volume which restrict the use of conventional, lower cost structure insulated with bonded tile thermal protection. The current solution is to use the CMC for control surface hot structure in areas which do not require their extreme high temperature properties. The result is high initial and recurring costs for these parts as well as weight penalties at high part counts. Without an order of magnitude reduction in thermal structure costs, commercial reusable access to space will be difficult, if not impossible, to achieve.
It is therefore a principal object of the present invention to provide a new construction for a ruddervator for an aerospacecraft which can be produced more inexpensively from a simpler fabrication process, and which has improved life and reliability over the conventional mechanically fastened upper skin-to-substructure construction presently in use for ruddervator applications.
It is another object of the present invention to provide a hybrid control surface for an aerospacecraft which can be manufactured more economically, which is simpler to repair, and which does not make use of typical mechanical fasteners to secure an upper skin to a substructure.
It is still another object of the present invention to provide a ruddervator for an aerospacecraft having a simplified design which requires significantly fewer independent component parts being needed for the construction of the ruddervator.
The above and other objects are provided by an airfoil for an aerospacecraft. The airfoil comprises a ruddervator having a oxide fiber/oxide matrix-based ceramic matrix composite (oxide-CMC) fabric which is secured to a ceramic foam insulation in the shape of an airfoil section when viewed cord-wise. A plurality of airfoil sections attached adjacent to one another form the ruddervator.
The oxide-CMC fabric is fused over the rigid ceramic foam insulation. The rigid foam insulation includes a hollowed out area through which at least one frame element extends. The hollowed out area of each airfoil section includes a plurality of integrally formed securing members, which in the preferred embodiment comprise lugs, which are secured to structure on the frame element. In one preferred form the frame element comprises a titanium box beam.
Each of the airfoil sections are secured to the frame element such that a lower end of one panel is positioned nestably within an upper end of its lower adjacent airfoil section. The frame element is secured to a torque box of the aerospacecraft such that the entire airfoil can be rotated as needed during flight.
The airfoil of the present invention thus does not require mechanical fasteners to be used to secure a skin to an independent substructure. The construction of the present invention further serves to reduce the cost and weight of the airfoil in large part because of the lower cost, higher specific strength and stiffness of the materials employed. Cost is also reduced because with the common design of the nesting airfoil sections, a single female lay-up mold can be used for the fabrication of all of the oxide-CMC fabric/ceramic foam insulation airfoil sections. The manufacturing cost is further reduced by utilizing the reduced tooling complexities of oxide-CMC fabrication processes over CMC fabrication processes.